Competition within the commercial and military aviation industries has compelled airframe manufacturers to develop aircraft that operate over ever-increasing flight envelopes. Modern advances in turbofan engine design and development have strived to improve the operating efficiencies of aircraft powerplants to keep abreast of these demands. One such advancement has been the implementation of sophisticated technologies and concepts in the aerodynamic design of fan or compressor blades for turbofan engines. However, in some instances, the engine manufacturer in providing a design to satisfy the airframe manufacturer's demands inadvertently generates an unfavorable aerodynamic characteristic, which impacts the operation of the system within its design envelope. On such occasions, the engine may experience operating difficulties that limit the performance of the aircraft, and in some cases, impact the structural integrity of the engine and/or its components.
As fan blade designs become more highly evolved to accommodate larger operating envelopes, the frequency at which these flow problems are encountered will increase. The scope and characteristic of these disturbances are varied and can include local flow separation on the fan blade or in some cases the establishment of unsteady aerodynamic loadings propagated by unfavorable boundary layer and shock wave development. On occasion, such fluctuations can match the resonant frequencies of the blade and disc assembly and generate further amplification of the detrimental effects at the fan blade and the associated stresses that are induced, particularly at the interconnection between the blade and disc.
These characteristics may occur at the corners of the operating envelope, or in rare cases may develop at or near the principal operating point for the engine and/or airframe. Prolonged exposure to these undesirable conditions may produce structural degradation of the fan blade and/or eventual catastrophic failure of the assembly. Without a viable solution to correct these characteristics the manufacturer is forced to impose operational and maintenance limitations on the turbofan assembly, which limits the aircraft's operating envelope, and the useful life and durability of the engine(s), and in some cases the airworthiness of the airframe. These limitations precipitate unforeseen costs for the operators through increased down time, elevated inspection periods and higher maintenance costs.
One purpose of the invention disclosed herein is to stabilize the boundary layer flow on the engine fan or compressor blades to mitigate fluctuating peak stresses sometimes experienced in the fan blade and fan disc under normal operations. The invention is a unique application of aerodynamic boundary layer control devices to this specialized flow problem in engine fan or compressor blade design.
Flow passing adjacent to the surface of a turbofan fan blade is dominated by the viscous characteristics of the air. This region of airflow is referred to as the boundary layer, and is characterized as either laminar or turbulent. The chordwise extent of the laminar and turbulent flow on the fan blade will depend upon atmospheric conditions, the operation of the engine and the geometry of the blade. A region of laminar flow will normally develop at the leading edge of the fan blade. Some distance downstream, the boundary layer will transition from laminar to turbulent flow. The chordwise location of the transition point will often vary along the length of the blade between its root and tip (see 23 in FIG. 1). Under normal conditions this transition point which identifies the end of laminar flow and the start of turbulent flow will occur at a nominally static point between the leading and trailing edge of the fan blade.
The operating efficiency of the fan or compressor blade may be greatly enhanced if its aerodynamic drag is maintained at a low level. One method to accommodate this goal is to develop and sustain a laminar boundary layer over the surface of the fan blade. Recent advancements in modern fan blade design have seen the development of airfoil geometries characterized by increasingly large regions of chordwise laminar flow. In such a laminar boundary layer, the air moves downstream over the blade following a relatively smooth trajectory and without appreciable mixing between different layers of the air stream. It is also known that certain airfoil configurations can generate or promote laminar flow. These types of airfoils are sometimes referred to as low drag or “laminar flow” airfoils, and are designed specifically for this purpose.
Other airfoil designs, such as those found in modern fan or compressor blades, may experience large regions of laminar flow simply because of their characteristic dimensions and/or the conditions under which they operate. In such examples, the thickness of the fan blade relative to its chord length is small and the blades operate at combinations of airspeed, altitude and ambient temperature where conditions are favorable for the development of a laminar boundary layer over a large percentage of the fan blade chord. Under these conditions the thickness of the laminar boundary layer will increase along the chord of the fan blade as air is entrained into the boundary layer. The principal mechanism responsible for the entrainment of additional air into the boundary layer is the viscous or shearing forces existing between adjacent layers of air. When the laminar boundary layer thickens, however, it becomes unstable and extremely sensitive to slight flow disturbances or geometric differences (in some cases caused by manufacturing tolerances) in the associated fan blade which can promote premature transition to a turbulent boundary layer. The chordwise location on the fan blade at which this transition occurs is referred to as the transition point, the position of which becomes relatively unstable the further the laminar boundary layer extends from the leading edge of the fan blade rearward along its chord.
Downstream of the transition point, the flow is turbulent and is characterized by an irregular or random motion superimposed upon the average or mean downstream flow and involves significant mixing of mass and momentum between layers of air within the boundary layer. This randomized mixing generates an exchange of momentum between the lower speed flow within the boundary layer and the higher speed flow outside the boundary layer. Entrainment of air into the turbulent boundary layer, generated by turbulent mixing, causes the boundary layer to thicken much faster than the more steady state laminar boundary layer. The increased thickness of the turbulent boundary layer alters the flow characteristics on adjacent fan blades and in the passage between these blades. These changes produce variations in most flow field parameters, including static pressure, velocity and mass flow.
In the more common case of compressible flow, a shock wave will often develop in the passage between adjacent fan blades. Under these conditions the thickness of the boundary layer has a direct influence on the chordwise position of the shock wave. On occasion, any chordwise oscillatory motion of the transition point will produce significant and unsteady changes in the position of the shock wave on the chord of the fan blade. Whenever the transition point is located upstream, or ahead of the shock wave, the chordwise position of the shock wave will move in response to any unsteadiness in the thickness of the turbulent boundary layer propagated by instabilities in the laminar to turbulent boundary layer transition point. In other applications, the chordwise movement of the shock wave will be directly affected by the movement of the transition point whenever the shock wave is located at the same chordwise position as the transition point, or in its immediate vicinity.
Promoting large regions of laminar flow will produce a higher operating efficiency for the powerplant and increased performance for the aircraft. However, a large region of laminar flow is often terminated by an unstable transition to a turbulent boundary layer which is characterized by an erratic and/or unsteady streamwise movement of the transition point along the chord of the fan blade. Fluctuations in the position of the transition point will impact the thickness of the turbulent boundary layer that forms immediately downstream of this location. Changes in the thickness of the turbulent boundary layer along the fan blade surface influence the flow characteristics on and between adjacent blades. Any unsteadiness in these flow characteristics will produce fluctuations in the aerodynamic loads on the fan blades, which in turn expose the blade and disc material to cyclic stresses with the potential to produce structural degradation, cracking and ultimately failure of the fan assembly.
Traditional methods utilized to resolve flow disturbances on engine fan blades include the application of grooves in the engine inlet casing and/or the cropping and grooving of the fan blade tips in an attempt to dampen the unsteady aerodynamic loading on the fan blades and disc. However, these techniques do not address the source of the aerodynamic problem identified herein, and provide only limited success in resolving the fundamental problem for which the invention described herein has been developed.
A primary goal of the present invention is to suppress any oscillations in the aerodynamic load on the fan blade and any tendency for subsequent structural resonance. This invention achieves this purpose by stabilizing the unsteady transition between a laminar and turbulent boundary layer located on the surface of the fan blade through the application of aerodynamic control devices to resolve this unique problem. The success of the present invention to reduce and stabilize the fluctuating fan blade and disc stresses is achieved by the strategic placement of the invention on the turbofan fan blade.